The present invention relates to improvements made to gas turbine blades for an airplane engine. More particularly, the invention relates to the cooling circuits of such blades.
It is known that the moving blades of an airplane engine gas turbine, and in particular of the high pressure turbine, are subjected to very high temperatures from combustion gases when the engine is in operation. These temperatures reach values that are well above those that can be withstood without damage by the various pieces that come into contact with these gases, thereby limiting the lifetime of such pieces.
Furthermore, it is known that raising the temperature of the gases in a high pressure turbine improves the efficiency of an engine, and thus the ratio of engine thrust over the weight of an airplane propelled by the engine. Consequently, efforts are made so as to provide turbine blades that are capable of withstanding higher and higher temperatures.
In order to solve this problem, it is known to provide such blades with cooling circuits seeking to reduce the temperature of the blades. By means of such circuits, cooling air, generally inserted into the blade via its root, passes through the blade following a path defined by cavities formed inside the blade prior to being ejected via orifices opening through the surface of the blade.
By way of example, a method is known which consists in providing the blade with a central cavity fed with cooling air from the root of the blade so as to xe2x80x9cimpactxe2x80x9d the leading edge thereof. However, since said cavity is separated from the hot gases solely by the wall of the blade, the air passing along the cavity heats up as it approaches the leading edge.
U.S. Pat. No. 5,720,431 discloses a blade provided with a central cavity surrounded by radial cavities situated on the concave and convex sides of the blade. That document also provides communication orifices enabling cooling air to pass between the central cavity and the radial cavities so that the air which passes through the radial cavities heats up little and loses little of its effectiveness in protecting the leading edge against hot gases.
The present invention thus seeks to mitigate those drawbacks by proposing improvements to gas turbine blades, and more particularly to the cooling circuits thereof, so as to limit the temperature of the leading edges of such blades. This has the consequence of increasing the lifetime of such blades.
To this end, the invention provides a gas turbine blade for an airplane engine, the blade comprising at least a first cooling circuit comprising at least a concave side cavity extending radially beside the concave face of the blade, at least a second cooling circuit independent of the first cooling circuit comprising at least one convex side cavity extending radially beside the convex face of the blade, and at least one third cooling circuit independent of the first and second cooling circuits and comprising at least one central cavity situated in the central portion of the blade between the concave side cavity and the convex side cavity, at least one leading edge cavity situated in the vicinity of the leading edge of the blade, communication orifices opening out into the central cavity and into the leading edge cavity, and outlet orifices opening out into the leading edge cavity and through the leading edge of the blade.
The presence of cavities surrounding the central cavity of the blade and the use of independent cooling circuits for the various cavities enables the blade to be protected in substantially uniform manner all along its height. As a result, the leading edge of the blade benefits from lower temperature cooling air.
In addition, in the central portion of the blade, the cooling of the convex and concave faces of the blade is provided by two completely independent circuits, thereby making it possible to control the concave and convex side temperatures of the blade independently by controlling the flow rate of cooling air traveling in each of the two circuits.
In addition, there is no radial air circulation in the cavity situated in the vicinity of the leading edge. The cooling air is exhausted directly into the combustion stream via the outlet orifices positioned in the leading edge of the blade, thus serving to avoid disturbing the effect of the impact jets by a shearing flow.